With reference to FIG. 1, a ducted fan gas turbine engine is generally indicated at 10 and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
Each of the compressors 13, 14 typically has a row of inlet guide vanes and a plurality of compressor stages, each stage comprising a set of stator vanes which receive and redirect the working fluid issuing from the rotating blades of the preceding stage. As aero engines have to operate at varying speeds and inlet conditions, it can be advantageous to be able to alter the aerodynamic flow angle of individual inlet guide vanes and stator vanes within the gas turbine annulus, depending upon the present engine operating speed and conditions. Vanes whose flow angles are alterable in this way are known as variable vanes.
A large variety of systems are conventionally used to actuate variable vanes. For example, unison rings can be used to rotate variable vanes about their radial axes and thereby change the aerodynamic flow angle. Each unison ring encircles the engine and is rotated by one or more actuators to produce movement in the circumferential direction. This movement can be converted by an arrangement of levers and spindles into the rotation of the variable vanes.
When developing new variable vane actuation systems it can be desirable to determine levels of system hysteresis and movement accuracy. In this context, by “hysteresis” is meant the amount of actuator movement prior to producing an actual vane angular movement. The hysteresis is typically caused by system mechanical wind up generated via a combination of aerodynamic and frictional loading and also mechanical backlash at articulation joints.
One known approach for determining system hysteresis is to measure vane and actuator relative movements via potentiometers attached to certain variable vanes within the stage and to compare those measurements with that of the output of a linear transducer(s) positioned and attached to the actuator(s).
However, conventionally it is necessary to perform such determinations as part of a full engine test. This is a disadvantage in that, during engine development, access to engine test facilities may be limited and expensive.